Technical Field
The present invention relates to gas turbine engine blade platforms.
Background Information
Gas turbine engines blades found in a high pressure compressor typically include a blade platform integrally disposed between an airfoil and a blade root for mounting the blade in a slot of a compressor disk. Under certain rotor speed and engine operating conditions, the excitation of certain modes of the blades from unsteady-state air flow on the blade platform can cause the blade platform corner to lose material. Consequentially, it can cause escalating downstream damage. However, complete redesign of a compressor blade for an existing matured engine is very costly and time-consuming as well as not being easy, and in some cases, impossible to retro-fit.
Thus, it is desirable to have a platform redesign and method of retrofitting blades with blades having redesigned platforms that avoid excitation of certain modes of the blades from unsteady-state air flow on the blade platform act to avoid the blade platform corner losing material.